The present invention relates to gas turbine engines and, more particularly, to apparatus for cooling a transition duct employed to conduct hot gasses from a combustor to a turbine stage of an advanced heavy duty gas turbine engine.
A large heavy duty gas turbine engine conventionally employs a plurality of cylindrical combustor stages operated in parallel to produce hot energetic gas for introduction into the first turbine stage of the engine. The first turbine stage preferably receives the hot gas in the shape of an annulus. A transition duct is disposed between each of the combustor stages and the first turbine stage to change the gas flow field exiting each combustor from a generally cylindrical shape to one which forms part of an annulus. The gas flow from all of the transition ducts thus produces the desired annular flow.
As well known, the thermodynamic efficiency of which a heat engine is capable depends on the maximum temperature of its working fluid which, in the case of a gas turbine, is the hot gas exiting the combustor stages. The maximum feasible temperature of the hot gas is limited by the operating temperature limit of the metal parts in contact with this hot gas, and on the ability to cool these parts below the hot gas temperature. The task of cooling the transition duct of an advanced heavy duty gas turbine engine, which is the one addressed by the present invention, is difficult because currently known cooling methods are either inadequate, or carry unacceptable penalties.
In a conventional heavy duty gas turbine engine, the entire external surface of the transition duct is exposed to relatively cool air discharged from the compressor, which supplies the total air flow for the gas turbine. The flow of air over the exterior of the transition duct to the combustor causes passive cooling. Some portions of the exterior of the transition duct are relatively well cooled by passive cooling, but others are poorly cooled thereby. Additionally, the portions of the exterior of the transition duct that are most poorly cooled are generally in structurally weaker areas, which are also areas more highly heated by the hot gas therewithin. To avoid failure resulting from excessive metal temperatures, the maximum combustor exit temperature must be limited by the maximum allowed metal temperature of the most poorly cooled areas of the transition duct. As heavy duty gas turbine combustor exit temperatures have been raised to promote increased thermal efficiency, various means to cool actively the relatively hot areas of the transitional duct have been employed. In an advanced heavy duty gas turbine, for which the combustor exit temperature is to be significantly higher than the approximately 2000 degrees usual to heavy duty gas turbines, the entire surface of the transition duct must be actively cooled, so that metal temperatures are kept to an acceptable level.
Known methods for cooling the walls of combustors permit air discharged by the compressor to pass through the combustor wall, and then direct it along the inside surface thereof, as a film to protect it from direct contact with the hot gas. This arrangement permits the combustor wall to operate significantly below the temperature of the hot gas. This film cooling method has been used for limited areas of the transition duct, especially those poorly cooled areas described above. However, the use of such film cooling is limited by the amount of air available exclusively for cooling the combustor and transition duct walls. This amount is typically less than thirty percent of the total air flow available to the combustor. For an advanced heavy duty gas turbine engine, virtually all of the air available for film cooling is required for cooling the combustor walls, and very little is available for cooling the transition duct walls. This limited availablity of cooling air flow comes about because approximately half of the total combustor air flow is required for complete combustion of the fuel and another quarter of the air flow is required for dilution and shaping of the hot gas profile exiting the combustor as required by the first turbine stage for acceptable efficiency and component life. These proportions can be altered slightly, depending on the particular design choices in a gas turbine engine, but a variety of practical obstacles blocks any large departure from them.
Another cooling technique which has found use in cooling the exterior of the transition duct employs an impingement plate, baffle or sleeve disposed a short distance away from the transition duct outer surface. The impingement sleeve contains an array of holes through which compressor discharge air passes to generate an array of air jets which impinge on and cool the outer surface of the transition duct.
U.S. Pat. No. 3,652,181 discloses such an impingement cooling technique for a transition duct in which the impingement sleeve surrounds only a portion of the transition duct. After impacting the surface to be cooled, the spent impingement air flows in the space between the transition duct outer surface and the impingement sleeve, towards holes in the transition duct. The air passing through these holes mixes with, and reduces the hot gas temperature just ahead of, the root area of the turbine blades and thus helps reduce the metal temperature of this portion of the turbine blades. Depending upon the heat transfer rate from the hot gas and the maximum allowed metal temperature, this method can use less cooling air than film cooling to maintain acceptable metal temperatures, and can be used in combination with film cooling to further reduce metal temperature. However, even the combination of impingement and film cooling for a transition duct would require more cooling air than is available in an advanced heavy duty gas turbine.
Further disclosure of impingement cooling of a gas turbine combustion component is found in U.S. Pat. No. 4,339,925. Although it is directed toward cooling a type of gas turbine combustion component which is completely different from that towards which the present invention is directed, this patent discloses typical elements of an impingement cooling system. There is disclosed therein a shell which has an array of holes through which cooling air passes to impinge on a hot gas casing towards the combustor. An embodiment is illustrated and described in which the impingement air flows along the hot gas casing eventually to enter the combustion process. A restrictor is disclosed for aiding the ejection of air from the space between the hot gas casing and the perforated shell. This patent recognizes that the number of inlet openings, as well as the spacing of the shell from the hot gas casing, represent variables which can be employed to produce the cooling effects required by the situation.
It can be seen from the prior art, as disclosed in U.S. Pat. Nos. 3,652,181 and 4,339,925, that impingement cooling of a combustion component can either consume a portion of the air flow allocated to the combustion process, or be performed in series with the combustor such that the air used to cool a combustion component is subsequently used in the combustion process. It is the series mode of cooling a transition duct which is addressed by the present invention.
For reasons which are well known by those skilled in the art of gas turbine design, there is a pressure drop or loss associated with forcing the compressor discharge air through openings in the combustor wall, to mix and burn with the fuel. This same pressure drop promotes the film cooling of the combustor and the dilution air jets which, in turn, shapes the temperature pattern of the air exiting the early portion of the combustor. Typically, this pressure drop falls between two and four percent of the compressor discharge pressure and, for reasons of thermal efficiency, is kept as low as possible. If the pressure drop is too low, poor mixing of the fuel and air, and resultant poor combustion, will result. If the pressure drop is too high, the gas turbine thermodynamic efficiency will be reduced.
In order to achieve impingement cooling, a pressure drop is required across the impingement sleeve or baffle, thereby forcing the cooling air through the holes at a sufficiently high velocity to achieve the required heat transfer rate. Generally, higher cooling rates are achieved by a higher pressure drop. Thus, it can be seen that employing impingement cooling of a transition duct in a series air-flow arrangement, will create an additional pressure drop to the combustion system which, if not kept to the lowest possible level, could cause a reduction in thermal efficiency greater than the increase obtained by raising the combustor exit temperature.
The pressure drop of an impingement cooling system essentially is generated by two components. First, there is a pressure drop needed to accelerate the air through the impingement sleeve holes to create the jets which impinge of the surface to be cooled. The second is more subtle, and is largely ignored in other known impingement cooling applications.
If the spent impingement air is to be used in the combustor, it must be collected and brought to the combustor. The collection naturally takes place between the impingement sleeve and the external surface of the transition duct, and it will be seen that, as one moves towards the combustor, the air flow velocity must steadily increase as more air is collected. The second component of pressure drop occurs due to the requirement to reaccelerate each additional quantity of spent impingement air to the velocity of that air already moving towards the combustor.
The local magnitude of the heat transfer in an impingement cooling system is determined by a number of variables. In particular, these variables include the cooling air properties, the local distance between the impingement sleeve and the transition duct surface, the hole size, spacing and array pattern, the impingement air jet velocity, and the velocity of air flowing perpendicular to the air jet such as, for example, air resulting from the collection of spent impingement air.
It can be seen that the number of variables which affect both the magnitude of heat transfer and the pressure drop of the overall impingement cooling system is large. The invention addresses the interplay of these variables for a successful design of a fully cooled transition duct for an advanced heavy gas turbine engine.
An air jet formed by an opening in an impingement plate must traverse the space separating the impingement plate from the surface to be cooled, and must impact the surface to be cooled with sufficient velocity and in sufficient volume to effect the desired cooling. The analysis of such jet impingement is relatively simple when only a single jet is involved. However, when an array of jets is used, the impingement air flowing away after impingement from one jet, captured between the surface being cooled and the impingement plate, tends to produce a crossflow of air which interfers with the cooling action of other jets, particularly those downstream in the direction along which the impingement air must flow to exit the constraining space. That is, a crossflow of air passing through the space between an aperture and the surface to be cooled may prevent the aperture-produced air jet from reaching the surface to be cooled, or may reduce the effectiveness of any portion of the air jet which may reach the surface to be cooled. The actual cooling effects of an array of jets is difficult to predict, and so may only be derived empirically.
The greater the velocity of the crossflow, the more the crossflow interferes with the effectiveness of the air jets. In the case of an impingement cooled transition duct in which all of the impingement air must flow outwardly from between the transition duct and the impingement plate, the amount of crossflowing air and its velocity increases systematically as it moves toward the exit. The increased velocity may partically or completely destroy the effectiveness of impingement jets located downstream thereof. It may be for this reason that a number of prior art devices employing impingement cooling of a transition duct (or a hot gas casing) provide for injecting the use impingement air into the interior of the transition duct. As discussed, this inefficient use of available cooling air is unacceptable for an advanced heavy duty gas turbine design.